Fabrication of composite laminates using temporarily stitched preforms

ABSTRACT

A composite structure is fabricated using a preform comprising a stack of unidirectional prepreg plies that are stitched together. During curing of the prepreg, the stitches melt and dissolve.

BACKGROUND INFORMATION 1. Field

The present disclosure generally relates to preforms used in thefabrication of composite laminate structures, and deals moreparticularly with a temporarily stitched preform.

2. Background

High performance composite structures may be fabricated by laying upplies of prepreg or by resin infusion of dry fibers. The fibers may bein unidirectional, woven or braided fabric form. In some applications,in order to reduce layup time, adjacent plies of the fabric may beco-stitched together using stitch material that remains with thecompleted structure after the structure is cured. Stitching the pliestogether allows the plies to the laid up on a tool in ply groups, ratherthan one-by-one, thereby increasing efficiency of the layup process.Co-stitched fabric plies formed from unidirectional reinforcement fibersare relatively formable, making them well-suited for forming highlycontoured structures, however structures fabricated using stitched pliesmay have less than the desired level of strength and crack resistance.

Accordingly, there is a need for a method of fabricating compositestructures using co-stitched plies of fabric which reduces or eliminatesthe presence of stitch material in the cured structure. There is also aneed for a preform used in the fabrication of such structures that canbe assembled using a co-stitched, multi-layer prepreg, or co-stitchedfiber layers suitable for resin infusion.

SUMMARY

The disclosed embodiments provide a method of fabricating a compositelaminate structure using a co-stitched multi-layer preform. In oneembodiment, the preform is formed by stitching together plies ofunidirectional prepreg using stitch material that melts during curing ofthe prepreg. In another embodiment, the preform is formed by stitchingtogether multiple fiber layers using stitch material that melts duringcuring of resin used to infuse the fiber layers. Melting of the stitchmaterial during the cure process effectively dissolves the stitches andavoids possible crimping between the stitches and the reinforcingfibers. Dissolution of the stitches reduces or eliminates stressconcentrations in the cured structure caused by fiber crimping, therebyimproving the mechanical performance of the composite laminate structurewhile reducing the possibility of crack propagation through thelaminate. The use of a co-stitched multi-ply preform may increaseproduction efficiency by allowing simultaneous layup and forming ofmultiple layers of fiber reinforcements.

According to one disclosed embodiment, a method is provided offabricating a composite structure. Prepreg plies are stitched togetherinto a stitched stack of prepreg plies having varying fiberorientations, and the stitched stack of prepreg plies is thermallycured. The stitching melts during thermal curing of the stitched stackof prepreg plies. The method may further comprise assembling the prepregplies into a stack, wherein each of the prepreg plies has resin tack,and assembling the prepreg plies into the stack includes using the resintack to adhere the plies together and maintain the fiber orientations ofthe prepreg plies during the stitching. Stitching the prepreg pliestogether is performed by using stitches that pass substantially throughthe thickness of the stack. Assembling the prepreg plies into a stackincludes laying down prepreg tows, and varying the fiber orientations ofthe tows for each of the plies. The method may also comprise debulking,consolidating and curing the stitched stack under a vacuum. Melting thestitching is performed before the stitched stack of prepreg plies isfully cured. The method may further comprise forming the stitched stackof prepreg plies into a desired shape corresponding to the shape of thecomposite structure.

According to another embodiment, a method is provided of making acomposite preform. A stack of prepreg plies is assembled, wherein eachof the plies includes reinforcing fibers held in a thermally curableresin matrix. The prepreg plies are stitched together after the stackhas been assembled. The stitching is performed using stitching materialthat melts during thermal curing of the prepreg plies. Assembling thestack of prepreg plies includes maintaining the plies in registrationrelative to each other by tacking the plies together. Tacking the pliestogether is performed using tack of the resin matrix in each of theplies. Assembling the stack of prepreg plies includes using the resinmatrix in each of the prepreg plies to hold the reinforcing fibers inthe plies in spaced relationship to each other during the stitching. Thestitching includes placing stitches substantially completely through thethickness of the stack of prepreg plies. The stack may be assembled bylaying prepreg tows, and the stitching may be carried out by placingstitches between the prepreg tows that pass substantially completelythrough the stack of prepreg plies. During assembly of the stack, theplies are oriented such that they have differing fiber orientations.

According to still another embodiment, a composite preform is provided.The preform comprises a stack of unidirectional prepreg plies havingvarying fiber orientations. Stitches passing through all of the prepregplies in the stack hold the plies together. The stitches are formed of astitching material capable of melting during thermal curing of theprepreg plies. Each of the prepreg plies includes prepreg tows, and thestitches pass between the prepreg tows. The stitches may be distributedgenerally uniformly across the stack of unidirectional prepreg plies.Each of the prepreg plies includes a resin matrix, and the stitchingmaterial is compatible with the resin matrix. The resin matrix may be athermoset resin, and the stitching material may be a thermoplasticresin. The thermoset resin has a cure temperature at which the thermosetresin is fully cured, and the thermoplastic resin has a melt temperaturethat is below the cure temperature of the thermoset resin.

According to still another embodiment, a method is provided offabricating a composite structure. Dry fiber plies are stitched togetherinto a stitched stack of dry fiber plies having varying fiberorientations. The stack of dry fiber plies is infused with a polymerresin. The resin infused stack is thermally cured. The stitching meltsduring thermal curing of the stitched stack. The method may also includedebulking, consolidating and curing the stitched stack under a vacuum.Stitching the dry fiber plies together is performed using stitches thatpass substantially through the entire thickness of the stack. The methodmay also comprise applying a tackifier to each of the dry fiber plies,and assembling the dry fiber plies into a stack, including using thetackifier to adhere the dry fiber plies together and maintain the fiberorientations of the dry fiber plies during the stitching. Assembling thedry fiber plies into a stack includes laying down dry fiber tows, andvarying the fiber orientations of the dry fiber tows for each of the dryfiber plies. The method may further comprise forming the stitched stackof dry fiber plies into a desired shape corresponding to the shape ofthe composite structure. The forming may be performed by forming thestitched stack of dry fiber plies onto a tool such as a mold.

According to another embodiment, a method is provided of making a dryfiber preform. The method comprises assembling a stack of dry fiberplies, each of which includes unidirectional reinforcing fibers. The dryfiber plies are stitched together after the stack has been assembled.The stitching is performed using stitches that pass through the stackand melt when they have been heated to a predetermined temperature. Themethod may also include applying a tackifier to each of the dry fiberplies, and assembling the stack of dry fiber plies includes using thetackifier to maintain the plies in registration relative to each other.

The features, functions, and advantages can be achieved independently invarious embodiments of the present disclosure or may be combined in yetother embodiments in which further details can be seen with reference tothe following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the illustrativeembodiments are set forth in the appended claims. The illustrativeembodiments, however, as well as a preferred mode of use, furtherobjectives and advantages thereof, will best be understood by referenceto the following detailed description of an illustrative embodiment ofthe present disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is an illustration of a top perspective view of a stitchedpreform.

FIG. 2 is an illustration of a bottom perspective view of the stitchedpreform shown in FIG. 1.

FIG. 3 is an illustration of a perspective view of a curved framesection that may be fabricated using a stitched preform.

FIG. 4 is an illustration of an exploded, perspective view of thestitched preform of FIG. 1, showing the individual layers of the preformand their respective fiber orientations.

FIG. 5 is an illustration of a plan view of the area designated as “FIG.5” in FIG. 1.

FIG. 6 is an illustration of a sectional view taken along the line 6-6in FIG. 5.

FIG. 7 is an illustration of a sectional view showing a portion of astitch between two tows at an early stage of curing.

FIG. 8 is an illustration similar to FIG. 7, but showing the stitchhaving melted into the surrounding matrix resin during a later stage ofthe curing.

FIG. 9 is illustration of a flow diagram of a method of fabricating acomposite structure using a stitched prepreg preform.

FIG. 10 is an illustration of a flow diagram of a method of fabricatinga stitched prepreg preform.

FIG. 11 is an illustration of a flow diagram of an alternate method offabricating a composite structure using a stitched, dry fiber preform.

FIG. 12 is an illustration of a flow diagram of aircraft production andservice methodology.

FIG. 13 is an illustration of a block diagram of an aircraft.

DETAILED DESCRIPTION

Referring to FIGS. 1 and 2, a composite preform 20 comprises a stitchedstack 22 of prepreg plies 24 a, 24 b, 24 c each of which hasunidirectional reinforcement in the form of fiber tows 28. The prepregplies 24 a, 24 b, 24 c in the stack 22, sometimes also referred toherein as “layers”, are tacked together by stitches 26 that extendthrough the thickness of the stack 22. Only the top and bottom of thestitches 26 are shown, respectively, in FIGS. 1 and 2. The preform 20may be used to fabricate any of a variety of composite structures,particularly those having simple or compound contours. For example,referring to FIG. 3, the preform 20 may be employed to fabricate aunitary composite frame section 30 by forming the stack 22 of prepregplies 24 a, 24 b, 24 c using suitable tooling (not shown), either beforeor after the prepreg plies 24 a, 24 b, 24 c are stitched together. Inthis example, the frame section 30 is curved along its length andcomprises a curved inner chord flange 34, a curved outer chord flange 36and a web 32. The flanges 34, 36 transition into the web along radiuscorners 38, 40 which have compounded curvatures. The frame section 30 ismerely illustrative of a wide range of composite laminate structuresthat may be fabricated using the disclosed preform 20. The frame section30 shown in FIG. 3 has a Z-shaped cross section, however othercross-sectional shapes are possible.

Referring particularly to FIGS. 1, 2 and 4, although three plies 24 a,24 b, 24 c are shown in the exemplary embodiment, the stack 22 maycomprise as few as two or greater than three plies 24, depending uponthe application. In the embodiment illustrated in FIGS. 1, 2 and 4, eachof the plies 24 a, 24 b, 24 c comprises a plurality of unidirectionalprepreg tows 28 that may be placed in multiple, side-by-side bandwidths(not shown) by automatic fiber placement equipment (not shown) or byother techniques. However, as will be discussed later, the stitchedstack 22 may comprise a stitched stack of dry fiber layers 24 a, 24 b,24 c of unidirectional dry fibers that may be in the form of tows,unidirectional tape, cut patterns of unidirectional reinforcement orother forms.

The prepreg tows 28 each comprise a bundle of individual reinforcingfibers (not shown) that is pre-impregnated with a suitable resin whichwill be discussed later in more detail. Each of the plies 24 a, 24 b, 24c may have any desired fiber orientation, but in the illustrated exampleshown in FIG. 4, respectively have 0°, 90° and 0° fiber orientations. Inone embodiment, the prepreg tows 28 may have a generally circularcross-sectional shape (see FIG. 6), while in another embodiment, theprepreg tows 28 may have a generally flat cross-sectional shape (notshown), sometimes referred to as a “flat tow” or a “spread tow”.

The resin used to impregnate the tows 28 may comprise a thermallycurable resin that is suitable for the application and has a desiredcure temperature. For example and without limitation, the reinforcingfibers may comprise carbon and the resin used as the matrix may comprisea thermally curable thermoset resin such as epoxy. Other types ofreinforcing fibers are possible, such as without limitation, metal,ceramic and/or glass fibers. Other types of resins may be employed asthe matrix, depending upon the application, such as, without limitationpolyester resins, vinyl ester resins, phenolic resins, polyimide resins,PBI (polybenzimidazole) resins, and BMI (bismaleimide) resins.

The presence of resin impregnated into the tows 28 causes the tows 28,and thus the plies 24 a, 24 b, 24 c to have resin tack, and this resintack causes the plies 24 a, 24 b, 24 c to adhere to each other when theyare laid up on top of each other. The adhesion provided by the resintack holds the plies 24 a, 24 b, 24 c in registration with each otherand in their desired ply orientations during subsequent processingdiscussed below in more detail. The matrix resin also holds the tows 28of the plies 24 in spaced relationship to each other through thethickness “t” of the stack 22. In some applications, it may be necessaryor desirable to apply a tackifier to the plies 24 a, 24 b, 24 c toincrease the adhesion between the plies 24 a, 24 b, 24 c. Similarly,where the tows 28 are dry (not impregnated with resin), a tackifier,sometimes referred to as a binder, may be used to adhere the layers 24a. 24 b, 24 c together and maintain their respective fiber orientationsuntil the stitched stack 22 can be formed into a desired shape.

The stitches 26 pass between the tows 28 and hold the plies 24 a, 24 b,24 c in their desired ply orientations. The number, density, size,spacing and type of the stitches used will depend upon the application.Similarly, the tightness of the stitches 26 may vary, depending upon thenumber of plies 24 in the stack 22 and the complexity of the compositestructure being fabricated. For example, where the composite structureis highly contoured, it may be desirable to employ relatively loosestitches 26 in order to allow the plies 24 a, 24 b, 24 c to slipslightly in-plane relative to each other as they are being formed overtooling. Slight in-plane slippage between the plies 24 a, 24 b, 24 c mayallow the stack 22 to better conform to contoured tool surfaces andavoid ply wrinkling and/bunching.

Referring now particularly to FIGS. 5 and 6, any of various types ofstitches 26 may be employed to stitch the plies 24 a, 24 b, 24 ctogether provided that the stitches 26 pass through substantially theentire thickness “t” (FIG. 1) of the stack 22, between any adjacent tows28 in each of the plies 24 a, 24 b, 24 c. In the illustrated embodiment,the stitches 26 are effectively looped around the tows 28, and extenddiagonally across the stack 22. However, in other embodiments, thestitches 26 may not be looped around all of the tows 28, and may extendin any direction across the stack 22. The stitches 26 may be formed andspaced apart from each other in any of a variety of manners, providingthat they adequately hold the plies 24 a, 24 b, 24 c together as thestack 22 is being formed over tooling (not shown) employed to shape thestack 22 into the desired shape of the composite structure. In someembodiments, however, it may be possible to stitch the plies 24 a, 24 b,24 c together after the stack 22 has been formed into a desired shape.

The material from which the stitches 26 is formed (hereinafter “stitchmaterial”) may comprise any of a variety of polymer resins that iscompatible with the matrix resin of the tows 28, and which has a melttemperature that results in melting of the stitches 26 during thermalcuring of the matrix resin. For example, the stitch material maycomprise a thermoplastic resin such as, without limitation, PEI(polyetherimide) PPS (polyphenylene sulphide), PES (polyethersulfone),PEEK (polyetheretherketone), PEKK (polyetheretherketone), and PEKK-FC(polyetherketoneketone-fc grade), which has a relatively low melttemperature that is within the range of temperatures required to curethe matrix resin. For example, where the matrix resin is an epoxy thatcures at approximately 180° C., the stitch material may comprise athermoplastic resin having a low melt temperature in the range of 150°C. In this example, the thermoplastic resin melts and combines with theflowable thermoset resin before the thermoset resin begins tosubstantially cure and harden. In one embodiment, a thermoplastic stitchmaterial is selected which remains intact to provide the necessarysupport of the plies 24 a, 24 b, 24 c, 24 d as the matrix resin meltsand initially becomes flowable. The thermoplastic stitch material maybegin to melt and dissolve into the matrix resin 44 only after theviscosity of the matrix resin 44 begins to increase as the matrix resin44 begins to harden during its initial stage of curing. Consolidation ofthe composite laminate structure is accomplished under vacuum which isused to debulk the plies 24 a, 24 b, 24 c and hold the plies 24 a, 24 b,24 c together without movement while the stitches 26 melt into the resinand the structure cures-consolidates.

FIG. 7 illustrates a cross-sectional side view of one of the stitches 26during an early stage of a cure cycle in which the formed compositelaminate structure is cured and consolidated by subjecting it to heatand pressure applied by a vacuum bag and/or an autoclave. Thecombination of applied heat and pressure causes the matrix resin 44 tobegin to flow, and consolidate the plies 24 a, 24 b, 24 c. The resinflow comes from the matrix resin 44 that is impregnated into the tows28. At this point in the cure cycle, the stitches 26 have not yet beenheated to their melt temperature, and therefore remain intact. As thetemperature is further increased during the cure cycle however, thestitch material begins to melt and flow 42 into the surrounding matrixresin 44 which is still flowable, until, as shown in FIG. 8, the stitchmaterial is fully dissolved within regions 46 of the matrix resin 44.The applied pressure aids in causing the stitch material and the matrixresin 44 to flow together and mix with each other. Depending upon theparticular polymer resin selected for use as stitches 26, the dissolvedstitch material may assist in toughening the matrix resin 44, and mayincrease mechanical properties, such as impact resistance, of the curedcomposite structure.

Attention is now directed to FIG. 9 which broadly illustrates theoverall steps of a method of fabricating a composite laminate structureusing a stitched prepreg. Beginning at step 48, a stack 22 ofunidirectional prepreg plies is assembled wherein plies may have varyingfiber orientations. Then at step 50, after the prepreg plies having beenassembled into a stack 22, the stack 22 may be formed into a desiredshape using tooling or other forming techniques. In some embodiments,however it may be possible to layup the preform 20 in a particular stackshape and then stitch the plies of the preform 20 together. In otherwords, the stitching of step 50 may be carried out after the forming ofstep 52. At step 54, the stitched and formed stack 22 is thermallycured, as by placing the stack into an oven or an autoclave. At 55,during thermal curing of the stitched stack 22, the stitching materialthat melts, causing the stitches to dissolve into the surrounding matrixresin 44 undergoing curing.

FIG. 10 broadly illustrates the overall steps of a method of making aprepreg preform 20 using prepreg plies that are stitched together withstitching material that melts during subsequent curing of the prepreg.At step 56, a stack 22 of prepreg plies is assembled. Each of the pliesincludes reinforcing fibers held in a thermally curable matrix resin 44.At step 58, the prepreg plies are stitched together after the stack hasbeen assembled, using a stitching material that melts and dissolvesduring thermal curing of the prepreg plies.

Referring again to FIGS. 1 and 2, as previously mentioned, in analternate embodiment, the preform 20 may be a dry fiber preform suitablefor use in any of various types of resin infusion processes in which thepreform serves as a reinforcement that is infused with resin. In thisembodiment, the preform 20 comprises a stitched stack 22 of the layers24 a, 24 b, 24 c, each of which is formed by a unidirectional dry fiberreinforcement such as fiber tows 28 (FIGS. 5 and 6) or unidirectionaldry fiber tape.

The layers 24 a, 24 b, 24 c have varying fiber orientations relative toeach other. The fiber tows 28 used in the dry fiber preform 20 maycomprise one or more materials similar to the materials discussed abovethat may be used to produce the fiber tows 28 of the prepreg embodimentof the preform 20. The dry fiber layers 24 a, 24 b, 24 c are temporarilystitched together by stitches 26 (FIGS. 5 and 6) that pass completelythrough the thickness “t” (FIG. 6). The stitches 26 hold the layers 24a, 24 b, 24 c together as a preform, but may be lose enough to allow thelayers 24 a, 24 b, 24 c to slip slightly relative to each other when thepreform 20 is formed down onto contoured surfaces of a tool (not shown)used in a resin infusion process. As previously mentioned, in someembodiments, the dry fiber layers 24 a, 24 b, 24 c may be formed into adesired shape before the dry fiber layers 24 a, 24 b, 24 c are stitchedtogether into a preform 20.

The stitches 26 assist in holding the layers 24 a, 24 b, 24 c in theirdesired orientations and in spaced apart relationship to each other asthe preform 20 is debulked, consolidated and infused with resin. Bymaintaining the dry fiber layers 24 a, 24 b, 24 c in their desiredorientations and spatial relationships until the matrix resin begins toharden with the onset of curing, the reinforcement of the curedcomposite structure may be more uniformly distributed and thereforecontribute to improving the mechanical performance of the compositestructure.

As in the previous prepreg preform 20 example, the material from whichthe stitches 26 are formed may comprise any of a variety of polymerresins that is compatible with the matrix resin used to resin infuse thepreform 20 after it has been placed on a tool. The stitch material has amelt temperature that results in melting of the stitches 26 duringthermal curing of the matrix resin following resin infusion of the dryfiber preform 20. For example, the stitch material used to stitch thelayers 24 a, 24 b, 24 c together as a dry fiber preform 20 may comprisea thermoplastic resin such as, without limitation, PEI (polyetherimide)PPS (polyphenylene sulphide), PES (polyethersulfone), PEEK(polyetheretherketone), PEKK (polyetheretherketone), and PEKK-FC(polyetherketoneketone-fc grade), which has a relatively low melttemperature that is within the range of temperatures required to curethe matrix resin used in a resin infusion process.

FIG. 11 broadly illustrates the steps of a method of fabricating acomposite structure using resin infusion of a dry fiber preform 20.Beginning at step 60, a stack of unidirectional dry fiber layers 24 a,24 b, 24 c is assembled, in which the layers have varying fiberorientations. At step 62, optionally, a tackifier may be applied to thelayers 24 a, 24, 24 c in order to assist in maintaining their respectivefiber orientations. At 64, the dry fiber layers 24 a, 24 b, 24 c arestitched together into a stitched stack 22, after the stack 22 has beenassembled. The stitches 26 hold the layers 24 a 24 b, 24 c of the stacktogether. At 66, the stitched stack 22 of dry fiber layers may be formedinto a desired preform shape.

Forming the stack 22 may be performed by forming the stack 22 ontotooling, either before or after the stack 22 has been stitched. Wherethe stack 22 is stitched before it is formed to a desired shape, and thetooling has one or more contours, the stitching 26 may allow the dryfiber layers 24 a 24 b, 24 c to slip slightly relative to each other inorder to better allow the layers to conform to contoured surfaces of thetool. Depending upon the type of resin infusion process being used, thedry fiber preform 20 may be transferred to a resin infusion tool at step68. In some embodiments, the tool on which the dry fiber layers 24 a, 24b, 24 c are formed into the shape of the preform 20 may be the tool thatis used during the resin infusion process. At step 70, the dry fiberpreform 20 is infused with resin, and at 72, the resin is thermallycured. The stitches 26 assist in holding the layers 24 a, 24 b, 24 c intheir desired orientations and in spaced apart relationship to eachother as the preform 20 is debulked, consolidated and infused withresin. At step 74, the stitching 26 that is used to hold the layers ofthe preform 20 together, melts and dissolve into the resin used toinfuse the preform 20.

Embodiments of the disclosure may find use in a variety of potentialapplications, particularly in the transportation industry, including forexample, aerospace, marine, automotive applications and otherapplication where composite laminate structures, particularly those thatare contoured and are fabricated in relatively high volume. Thus,referring now to FIGS. 12 and 13, embodiments of the disclosure may beused in the context of an aircraft manufacturing and service method 76as shown in FIG. 12 and an aircraft 78 as shown in FIG. 13. Aircraftapplications of the disclosed embodiments may include, for example,without limitation, composite laminate frame sections, spars, stringersand beams, to name only a few. During pre-production, exemplary method76 may include specification and design 80 of the aircraft 78 andmaterial procurement 82. During production, component and subassemblymanufacturing 84 and system integration 86 of the aircraft 78 takesplace. Thereafter, the aircraft 78 may go through certification anddelivery 88 in order to be placed in service 90. While in service by acustomer, the aircraft 76 is scheduled for routine maintenance andservice 92, which may also include modification, reconfiguration,refurbishment, and so on.

Each of the processes of method 76 may be performed or carried out by asystem integrator, a third party, and/or an operator (e.g., a customer).For the purposes of this description, a system integrator may includewithout limitation any number of aircraft manufacturers and major-systemsubcontractors; a third party may include without limitation any numberof vendors, subcontractors, and suppliers; and an operator may be anairline, leasing company, military entity, service organization, and soon.

As shown in FIG. 12, the aircraft 78 produced by exemplary method 76 mayinclude an airframe 94 with a plurality of systems 96 and an interior98. Examples of high-level systems 96 include one or more of apropulsion system 100, an electrical system 102, a hydraulic system 104and an environmental system 106. Any number of other systems may beincluded. Although an aerospace example is shown, the principles of thedisclosure may be applied to other industries, such as the marine andautomotive industries.

Systems and methods embodied herein may be employed during any one ormore of the stages of the production and service method 76. For example,components or subassemblies corresponding to production process 84 maybe fabricated or manufactured in a manner similar to components orsubassemblies produced while the aircraft 76 is in service. Also, one ormore apparatus embodiments, method embodiments, or a combination thereofmay be utilized during the production stages 84 and 86, for example, bysubstantially expediting assembly of or reducing the cost of an aircraft76. Similarly, one or more of apparatus embodiments, method embodiments,or a combination thereof may be utilized while the aircraft 76 is inservice, for example and without limitation, to maintenance and service92.

The description of the different illustrative embodiments has beenpresented for purposes of illustration and description, and is notintended to be exhaustive or limited to the embodiments in the formdisclosed. Many modifications and variations will be apparent to thoseof ordinary skill in the art. Further, different illustrativeembodiments may provide different advantages as compared to otherillustrative embodiments. The embodiment or embodiments selected arechosen and described in order to best explain the principles of theembodiments, the practical application, and to enable others of ordinaryskill in the art to understand the disclosure for various embodimentswith various modifications as are suited to the particular usecontemplated.

What is claimed is:
 1. A method of fabricating a composite structure,comprising: stitching composite layers together using thermoplasticstitches to form a stitched stack of composite layers having varyingfiber orientations, wherein stitching comprises extending thethermoplastic stitches diagonally through a thickness of the stitchedstack; and thermally curing the stitched stack of composite layers toform a cured composite structure, wherein thermally curing the stitchedstack of composite layers comprises: heating the stitched stack ofcomposite layers and applying pressure to the stitched stack ofcomposite layers to cause a matrix resin to become flowable; continuingto heat the stitched stack of composite layers and apply pressure to thestitched stack of composite layers to begin increasing a viscosity ofthe matrix resin; continuing to heat the stitched stack of compositelayers after the viscosity begins to increase; and melting the stitchingafter the viscosity of the matrix resin begins to increase, whereinmelting the stitching is performed before the stitched stack ofcomposite layers is fully cured so that the stitching and the matrixresin flow together and mix with each other in a region and thecomposite layers are held together without movement while the stitchesmelt into the matrix resin and into the composite structure.
 2. Themethod of claim 1, further comprising: assembling the composite layersinto a stack, wherein each of the composite layers has resin tack, andassembling the composite layers into the stack includes using the resintack to adhere the composite layers together and maintain the fiberorientations of the composite layers during the stitching.
 3. The methodof claim 1, wherein the stitched stack of composite layers has athickness, and stitching the composite layers together is performed byusing stitches that pass substantially through the thickness of thestitched stack; and wherein a thermoplastic material of each stitchcomprises a thermoplastic resin selected from a group consisting of PEI(polyetherimide) PPS (polyphenylene sulphide), PES (polyethersulfone),PEKK (polyetheretherketone), and PEKK-FC (polyetherketoneketone-fcgrade).
 4. The method of claim 2, wherein assembling the compositelayers into a stack includes: laying down prepreg tows, and varying thefiber orientations of the tows for each of the composite layers.
 5. Themethod of claim 4, further comprising: debulking, consolidating andcuring the stitched stack of composite layers under a vacuum.
 6. Themethod of claim 1, further comprising: forming the composite layers intoa desired shape corresponding to the shape of the composite structure.7. The method of claim 6, wherein forming the composite layers isperformed before stitching the composite layers together.
 8. The methodof claim 6, wherein forming the composite layers is performed afterstitching the composite layers together.
 9. The method of claim 1,wherein each stitch comprises a thermoplastic material that has a melttemperature of about 150 degrees.
 10. A method of fabricating acomposite structure, comprising: assembling composite layers into astack; stitching, using a thermoplastic resin, the composite layerstogether into a stitched stack of composite layers having varying fiberorientations, the composite layers comprising a matrix resin, whereinthe matrix resin is a thermoset resin, wherein stitching comprisesstitching diagonally through a thickness of the stack; thermally curingthe stitched stack of composite layers; and melting the stitching duringthermal curing of the stitched stack of composite layers, whereinmelting of the stitching during thermal curing occurs after a viscosityof the matrix resin begins to increase and before the stitched stack ofcomposite layers is fully cured.
 11. The method of claim 1, wherein themelted stitches toughen the matrix resin of the composite structure andincrease an impact resistance of the cured composite structure.
 12. Themethod of claim 1, wherein stitching further comprises looping thestitches around tows of top and bottom composite layers of the stitchedstack.
 13. The method of claim 10, further comprising: assembling thecomposite layers into a stack, wherein each of the composite layers hasresin tack, and assembling the composite layers into the stack includesusing the resin tack to adhere the composite layers together andmaintain the fiber orientations of the composite layers during thestitching.
 14. The method of claim 10, wherein a thermoplastic materialof each stitch comprises a thermoplastic resin selected from a groupconsisting of PEI (polyetherimide) PPS (polyphenylene sulphide), PES(polyethersulfone), PEKK (polyetheretherketone), and PEKK-FC(polyetherketoneketone-fc grade).
 15. The method of claim 10, whereinthe stitched stack of composite layers has a thickness, and stitchingthe composite layers together is performed by using stitches that passsubstantially through the thickness of the stitched stack.
 16. Themethod of claim 15, wherein assembling the composite layers into a stackincludes: laying down prepreg tows, and varying the fiber orientationsof the tows for each of the composite layers.
 17. The method of claim16, further comprising: debulking, consolidating and curing the stitchedstack of composite layers under a vacuum.
 18. The method of claim 10,further comprising: forming the composite layers into a desired shapecorresponding to the shape of the composite structure.
 19. The method ofclaim 18, wherein forming the composite layers is performed beforestitching the composite layers together.
 20. The method of claim 18,wherein forming the composite layers is performed after stitching thecomposite layers together.
 21. The method of claim 10, wherein eachstitch comprises a thermoplastic material that has a melt temperature ofabout 150 degrees.
 22. The method of claim 10, wherein thermally curingthe stitched stack of composite layers forms a cured compositestructure, and wherein the melted stitches toughen the matrix resin ofthe composite structure and increase an impact resistance of the curedcomposite structure.
 23. The method of claim 10, wherein stitchingfurther comprises looping the stitches around tows of top and bottomcomposite layers of the stitched stack.